7.0 LUNAR MODULE PERFORMANCE
7.1 STRUCTURAL AND MECHANICAL SYSTEMS
The structural loads were within design values for all phases of the
mission based on guidance and control data, cabin pressure measurements,
command module acceleration data, photographs, and crew comments.
Translunar docking loads were higher than those of previous missions
because of a pitch misalignment angle of 11 degrees between the command and
service module and the lunar module/S-IVB prior to docking probe retraction
to the hard-docked configuration. The bending moment during translunar
docking was computed to be 425,000 inch-pounds which approaches the design
limit of 437,000 inch-pounds.
The sequence films from the onboard camera showed no evidence of large
structural oscillations during lunar touchdown, and crew comments agree
with this assessment. Landing on the lunar surface occurred with estimated
velocities of 6.8 ft/sec in the minus X direction, 1.2 ft/sec in the plus Y
direction, and 0.6 ft/sec in the plus Z direction. The descent rate at probe
contact was 0.5 ft/sec. Following probe contact, the descent engine was
shut down while the footpads were still about 1.6 feet above the surface,
resulting in the 6.8 ft/sec velocity at footpad contact. Computer
simulations indicate 1.0 inch of stroke in each primary strut except the
forward strut, for which a 3.0-inch stroke is estimated. The simulations
also indicate that the forward footpad was off the surface in the final
rest position. The crew stated that the forward footpad was loose and
rotated easily, confirming the computer results.
At touchdown, the lunar module was located partially inside a small
crater with the rim of the crater directly underneath the descent engine
skirt. The descent engine skirt buckled during landing. This is accounted
for in the touchdown dynamic analysis, and was expected as the skirt length
had been extended 10 inches over that of previous vehicles. This buckling
was noted by the crew and confirmed by photographs of the damaged skirt
( fig. 7-1 ).
The crew reported that there was a gap between the exit plane of
the skirt and the lunar surface, indicating that buckling was probably
caused by a buildup of pressure inside the nozzle due to proximity to the
lunar surface, and not due entirely to contact of the nozzle skirt with the
lunar surface. The crew also reported that the buckling seemed to be uniform
around the skirt periphery and that the exit plane height above the surface
The vehicle contact velocity and attitude data at touchdown show that
the landing was very stable in spite of the relatively high lunar surface
slope at the landing point. The plus-Z and plus-Y footpads contacted the
lunar surface nearly simultaneously, providing a nose-up pitch rate of 17
deg/sec and a roll rate to the left of 15 deg/sec. Final spacecraft
settling occurred 1.8 seconds later. The vehicle at-rest attitude, as
determined from the gimbal angles, was 6.9 degrees pitch up and 8.6 degrees
roll to the left, resulting in a vehicle tilt angle on the lunar surface of
approximately 11 degrees from the horizontal
( fig. 7-2).
The performance of the electrical power distribution system and
batteries was satisfactory. Descent battery management was performed as
planned, all power switchovers were accomplished as required, and parallel
operation of the descent and ascent batteries was within acceptable limits.
The d-c bus voltage was maintained above 28.9 volts, and the maximum observed
current was 74 amperes, during powered descent. Electrical power used during
the mission is given in section 7.9.6.
The steerable antenna exhibited random oscillation
characteristics identical to those experienced on previous missions, and
resulted in three instances of temporary loss of voice and data. Also at
random times, small oscillations were noted and were damped out. The
problems with the antenna are discussed in section 14.2.4.
The lunar module did not receive VHF transmissions from the command
module during the descent phase of the mission. The checklist erroneously
configured the command module to transmit on 296.8 MHz and the lunar
module to receive on 259.7 MHz.
With the exceptions noted above, all functions including voice, data,
and ranging of both the S-band and the VHF equipment operated satisfactorily
during all phases of the mission.
The landing radar acquisition of slant range and velocity was normal.
The acquired slant range of 42,000 feet increased to about 50,000 feet in
approximately 10 seconds. The indication of range increase may have been
caused by blockage from a lunar mountain at initial acquisition. Velocity
was acquired at an altitude of approximately 39 000 feet above the local
terrain. Landing radar outputs were affected at an altitude of about 30 feet
by moving dust and debris.
Rendezvous radar tracking operation during the rendezvous sequence
was nominal. A lunar module guidance computer program was used after lunar
orbit insertion to point the rendezvous radar antenna at the command and
service module, thus enabling acquisition at approximately 109 miles. Two
problems were noted during the mission and are as follows:
a. During the VHF ranging system/rendezvous radar comparison test
after undocking and separation, a range difference of 2400 feet existed
between the rendezvous radar and VHF ranging systems. This difference
represents high-frequency ranging-tone cycle slippage in the rendezvous
radar, probably caused by excessive phase shift. Range errors associated
with cycle slippage, due to insufficient heater operation, have occurred
during system checkout and have produced phase shifts. Downlink data at the
time of the problem indicates that the rendezvous radar transponder heater
was not in operation when the rendezvous radar checkout was first
attempted; therefore, it is assumed that the phase shift was caused by low
7.5 CONTROLS AND DISPLAYS
b. Acquisition with the rendezvous radar during ascent was
unsuccessful. The radar antenna was pre-positioned prior to lunar lift-off
to an approximate lunar-module guidance-computer designated position for
acquisition following insertion. In this position, acquisition would have
been accomplished when the command module came into the rendezvous radar
antenna field of view. A review of lift-off television data revealed
rendezvous radar antenna movement during the first 2
seconds of flight.
Analysis has also shown that expansion of the ascent engine plume, after
being deflected from the descent stage structure, exerts sufficient
pressure on the antenna to overcome gimbal friction and move the antenna.
Radar acquisition apparently was not accomplished because the radar
antenna moved. Rendezvous radar tracking during ascent is not required.
The controls and displays performed normally. The range/range-rate
tape meter glass was broken with about 20 percent of the glass missing;
however, the meter operated satisfactory during the flight. Section 14.2.8
contains a discussion of this anomaly.
7.6 GUIDANCE, NAVIGATION, AND CONTROL
Guidance, navigation, and control system performance was satisfactory
throughout the mission except for two anomalies. There was a simultaneous
indication of an abort guidance system warning light and master alarm on two
occasions (sec. 14.2.6), and no line-of-sight rate information was
displayed on the Commander's crosspointers during the rendezvous braking
phase (sec. 14.2-7). Neither anomaly affected overall systems performance.
The primary guidance system was activated at 98 hours 26 minutes, the
computer timing was synchronized to the command module computer, and the
lunar module platform was aligned to the command module platform. The crew
had difficulty seeing stars in the alignment optical telescope while
performing the docked realignment of the lunar module platform, but this is
normal because of reflected light from the command- module structure.
is a summary of all platform realignments of the primary guidance
system, both in orbit and on the lunar surface. The calculated gyro drift
rates compare well with the 1 sigma value of 2 meru and indicate good gyro
performance. Accelerometer performance was equally good although shifts in
the X- and Y-accelerometer bias of 0.39 and 0.46 cm/sec2, respectively,
were detected while on the lunar surface. The shift resulted from removing
and reapplying power to the inertial measurement unit and is not unusual.
is a summary of preflight inertial component calibration data.
After a nominal separation from the command module, the abort
guidance system was activated, initialized, and aligned to the primary
is a summary of preflight and inflight
performance of the abort guidance system accelerometers and gyros.
The powered descent to the lunar surface was initiated on time.
is a sequence of significant events during descent. A landing site update
to move the targeted landing point 853 meters (2800 feet) downrange was
made 95 seconds after ignition. The computer began accepting landing radar
updates and began adjusting its estimate of altitude upward by 4800 feet.
After enabling landing radar updates, the primary guidance altitude flattened
out for approximately 70 seconds
This resulted from the initial
estimate of the slope stored in the computer being 1 degree; whereas, the
true mare slope was zero. Convergence occurred rapidly once the lunar module
was over the Apennine foothills where the computer-stored slope agreed more
closely with the actual slope.
Figure 7-3 is a time history of altitude from
the primary and abort guidance systems. Data indicate that 18 separate
deflections of the hand controller were made for landing point redesignations
during the approach phase program. The total effect of the redesignations
moved the landing site coordinates 338 meters (1110 feet) uprange and 409
meters (1341 feet) to the north. Touchdown disturbances were nominal despite
the 11-degree slope upon which the landing occurred.
is a time history of
spacecraft rates and attitudes at touchdown.
Performance during ascent was nominal. For the first time,
accelerometer biases were updated while on the lunar surface to correct for
the small expected shifts experienced when the system was powered down.
Since the lunar surface bias determination technique had not been totally
proven, only half of the measured shift in the X accelerometer bias was
corrected. As a result, some bias error existed during ascent and
contributed about 2 ft/sec to the radial velocity error. Analysis is
continuing to determine the cause of the remainder of the radial velocity
error and possible causes will be discussed in a supplement to this report.
Because the primary guidance system radial velocity was greater than
that from the powered flight processor and the abort guidance system, the
velocity residuals at engine shutdown were trimmed using the abort
guidance system solutions.
After trimming velocity residuals, an abort guidance system warning
and master alarm occurred. They were reset by the crew and a computer self-
test was performed successfully. System performance was nominal before,
during, and after the warnings. See section 14.2.6 for further discussion of
this anomaly. No vernier adjust maneuver was required, and the direct
rendezvous was accomplished without incident.
is a summary of
rendezvous maneuver solutions.
The Commander reported that there were no line-of-sight rate data
displayed on his crosspointers at a separation distance of 1500 feet.
However, line-of-sight rates existed at this time because thrusting upward
and to the left was required to null the -line-of-sight rates. Also, the
Command Module Pilot verified the presence of line-of-sight rates. Section
14.2.7 contains a discussion of this anomaly.
After a successful docking, the lunar module was configured for the
deorbit maneuver and jettisoned. The velocity changes observed at jettison in
the X, Y, and Z axes were minus 1.24, minus 0.01, and minus 0.05 ft/sec,
respectively. This is equivalent to a 206 lb-sec impulse. For comparison, the
separation velocities observed at undocking prior to powered descent were
minus 0.18, minus 0.02, and minus 0.04 ft/sec in the X, Y, and Z axes,
respectively, or an impulse of 205 lb-see. The close agreement indicates the
tunnel was completely vented and the impulse was due entirely to the
separation springs. After jettison, the deorbit maneuver was accomplished and
performance was nominal.
7.7.1 Reaction Control System
The reaction control system performed satisfactorily throughout the
mission with no anomalies. Skillful use of the system by the crew accounted
for the propellant consumption being well below predicted levels. Section
7.9.3 contains a summary of the consumables usage during the mission.
7.7.2 Descent Propulsion System
Data analysis indicates that the descent propulsion system performed
nominally during powered descent. The total firing time was 739.2 seconds .
The propellant quantity gaging system indicated about 1055 Pounds of usable
propellant remained at engine shutdown or about 103 seconds of hover time. The
supercritical helium system operated nominally. The skirt of the engine was
buckled during landing (sec. 7.1). Section 7.9.1 contains a summary of the
descent propulsion system consumables usage during the mission.
7.7.3 Ascent Propulsion System
The ascent propulsion system performance during the lunar ascent
maneuver and the terminal phase initiation maneuver was satisfactory. The
total engine firing time for the two maneuvers was 433.6 seconds. The ascent
propulsion system consumables usage is summarized in section 7.9.2.
7.8 ENVIRONMENTAL CONTROL SYSTEM
The performance of the environmental control system was satisfactory
throughout the mission. The waste management system functioned as
expected; however, the urine receptacle valve was inadvertently left open
for about 6 hours during the first lunar sleep period. This resulted in the
loss of about 8 pounds of descent stage oxygen before the crew was awakened
to close the valve.
The overspeed of the water separator which occurred on Apollo 14 during
cabin-mode operation was not evident during this mission because of a
decrease in flow with the helmet and gloves off that resulted from a
reconfiguration of valves and hose connections. The only off-nominal
performance of the water separator occurred following the cabin
depressurization for the standup extravehicular activity when the speed
decreased, causing a master alarm (see sec. 14.2.2).
After the first extravehicular activity, a broken quick disconnect
between the water bacteria filter and the water drink gun resulted in
spillage of about 26 pounds of water into the cabin (see sec. 14.2.3).
The water was cleaned up by the crew before the second extravehicular
Fluctuations in water/glycol pump differential pressure were noted
following the cabin depressurizations for the standup extravehicular
activity and the second extravehicular activity (see sec. 14.2.1). Otherwise,
the heat transport system functioned normally.
On Apollo 15, the suits were removed and dried for more than 1 hour
by connecting the oxygen umbilicals to the suits and allowing gas to flow
through them. This was accomplished at the beginning of each rest period
following the first two extravehicular activities.
All lunar module consumables remained well within red-line limits.
7.9.1 Descent Propulsion System
Propellant.- The descent propulsion system propellant load quantities
shown in the following table were calculated from known volumes and weights
of offloaded propellants, temperatures, and densities prior to lift-off.
Supercritical helium.- The quantities of supercritical helium were
determined by computations using pressure measurements and the known
volume of the tank.
7.9.2 Ascent Propulsion System
Propellant.- The ascent propulsion system total propellant usage was
approximately as predicted. The loadings shown in the following table
were determined from measured densities prior to launch and from weights
of off-loaded propellants.
Helium. The quantities of ascent propulsion system helium were
determined by pressure measurements and the known volume of the tank.
7.9.3 Reaction Control System Propellant
The reaction control system propellant consumption was calculated
from telemetered helium tank pressure histories using the relationships
between pressure, volume, and temperature.
The actual quantities of oxygen loaded and consumed are shown in the
The actual water quantities loaded and consumed, shown in the following
are based on telemetered data.
7.9.6 Electrical Power
The total battery energy usage is given in the following